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公开(公告)号：US12025075B2
公开(公告)日：2024-07-02
申请号：US18026601
申请日：2021-09-13
发明人： Cheng Cheng , Haiqing Zhou , Gui Tian , Jingyu Xiong , Yeming Zeng , Guofeng Zhou , Hongbo Xu
摘要： A cryogenic engine for a space apparatus is provided. The cryogenic engine includes an injector body, a thrust chamber, and a spark plug, where the injector body is provided therein with an accommodating space; the spark plug is provided on one side of the injector body, and an electrode provided on the spark plug extends into the accommodating space; the thrust chamber is provided on the other side of the injector body and is communicated with the accommodating space; the injector body is provided with a combustion improver flow channel and a combustible agent flow channel; and the combustion improver flow channel and combustible agent flow channel are connected with the accommodating space.
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公开(公告)号：US20240125290A1
公开(公告)日：2024-04-18
申请号：US18380892
申请日：2023-10-17
申请人： SPARC Research LLC
发明人： Guy B. Spear
CPC分类号： F02K9/95 , F02K9/346 , F05D2260/99
摘要： Solid rocket motors or ramjets may be provided. Such devices may include a combustion chamber containing at least one fuel, a metal phase alloy film within the combustion chamber, where the metal phase alloy film coupled to the at least one fuel, and an activator operably coupled to the metal phase alloy film, the activator including an electrical activation system, a laser activation system, or an initiator activation system. This allows combined functional features on pulse motors to release inhibitors and ignite solid fuel/propellant. Fast reactions can be achieved by controlling the thickness of the alloys and using multiple layers to achieve the areal energy density. Foils can be rolled and inserted into the fuel/propellant bore. When rolled, it can easily follow the grain bore diameter changes with temperature and provide no additional mechanical loads to the grain.
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公开(公告)号：US11512668B2
公开(公告)日：2022-11-29
申请号：US17106089
申请日：2020-11-28
申请人： Raytheon Company
发明人： Frederick B. Koehler , Jacob A. Pinello-Benavides , Curtis S. Copeland , Isaiah M. McNeil , Paul Kadlec , Lauren E. Brunacini , Mark T. Langhenry
摘要： A rocket motor has an electrically operated propellant initiator for a propellant grain that includes an electrode arrangement configured to concentrate an electric field at an ignition electrode for igniting an electrically operated propellant. The rocket motor includes a combustion chamber containing at least one propellant grain and an electrically operated propellant initiator operatively coupled to the propellant grain to initiate combustion of the propellant grain. The electrically operated propellant initiator includes the electrically operated propellant and at least one pair of electrodes configured to ignite the electrically operated propellant. The pair of electrodes includes a ground plane electrode and an ignition electrode. When an electrical input is applied to the electrically operated propellant initiator, the electric field is concentrated at the ignition electrode to ignite the electrically operated propellant at the location where the ignition electrode is arranged.
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公开(公告)号：US11499505B2
公开(公告)日：2022-11-15
申请号：US16896271
申请日：2020-06-09
申请人： Raytheon Company
摘要： A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.
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公开(公告)号：US20220307449A1
公开(公告)日：2022-09-29
申请号：US17213081
申请日：2021-03-25
申请人： Goodrich Corporation
摘要： A method for generating electric power for a rocket system includes burning a primary solid propellant grain to create a primary high pressure gas for providing thrust to the rocket, opening a first valve to divert a portion of the high pressure gas to an auxiliary solid propellant grain for igniting the auxiliary solid propellant grain, wherein the auxiliary solid propellant grain is disposed in a housing separate from the primary solid propellant grain, and burning the auxiliary solid propellant grain to create an auxiliary high pressure gas for turning a turbine. The method further includes driving a generator with the turbine and generating an electric power with the generator.
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公开(公告)号：US11220979B1
公开(公告)日：2022-01-11
申请号：US17094716
申请日：2020-11-10
发明人： Aaron Davis , Scott Stegman
摘要： An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.
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公开(公告)号：US20210231082A1
公开(公告)日：2021-07-29
申请号：US17094733
申请日：2020-11-10
发明人： Aaron Davis , Scott Stegman
摘要： An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.
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公开(公告)号：US11060482B2
公开(公告)日：2021-07-13
申请号：US15750757
申请日：2016-09-13
发明人： Hyun Duck Kwak , Jin Han Kim , Dae Jin Kim , Chang Ho Choi
摘要： The present invention relates to a liquid rocket engine using a booster pump driven by an electric motor, and more particularly, to a liquid rocket engine using a booster pump driven by an electric motor in which a booster pump is installed between a propellant tank and a propellant pump so that a requirement for an inlet pressure of the propellant pump may be met even in a state in which an internal pressure of the propellant tank is reduced, resulting in reduced amount of a propellant and reduced weight of the propellant tank, and the electric motor configured to drive the booster pump may be efficiently cooled through an oxidant, a cooling line, and the like.
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公开(公告)号：US10968865B2
公开(公告)日：2021-04-06
申请号：US15597922
申请日：2017-05-17
申请人： Airbus DS GmbH
发明人： Ulrich Gotzig , Malte Wurdak , Joel Deck , Manuel Frey
摘要： A rocket propulsion system comprises a combustion chamber, an oxygen supply system, comprising an oxygen supply duct and being configured to supply oxygen to the combustion chamber, and a hydrogen supply system, comprising a hydrogen supply duct and being configured to supply hydrogen to the combustion chamber. An ignition unit of the propulsion system, to which at least portions of the oxygen and the hydrogen supplied to the combustion chamber can be supplied, is configured to initiate combustion of the oxygen-hydrogen mixture in the combustion chamber. The propulsion system further comprises a cooling duct extending along an inner surface of a combustion chamber wall and through which at least a portion of the oxygen supplied to the combustion chamber, at least a portion of the hydrogen supplied to the combustion chamber or a combustion gas mixture emerging from the ignition unit flows.
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公开(公告)号：US20190309707A1
公开(公告)日：2019-10-10
申请号：US16449608
申请日：2019-06-24
申请人： The Boeing Company
IPC分类号： F02K9/95 , H01T13/34 , H01T13/20 , F02K9/94 , H01T13/38 , H01T13/41 , F02C7/266 , F02C7/264 , H01T13/22 , H01T13/54
摘要： A spark plug having an insulating body defining a longitudinal axis and including a base portion and an obstruction portion, a first electrode including a proximal portion, a sheathed portion and a distal portion, the sheathed portion of the first electrode extending through the base portion of the insulating body, and a second electrode including a proximal portion, a sheathed portion and a distal portion, the sheathed portion of the second electrode extending through the base portion and the obstruction portion of the insulating body, wherein the obstruction portion axially extends beyond the distal portion of the first electrode.
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